1. Field of the Invention
The invention relates to film cooling of hot surfaces such as those found in hot aircraft gas turbine engine components and, particularly, to film cooling holes such as those found in combustor liners and turbine nozzle airfoils in gas turbine engines.
2. Description of Related Art
A typical gas turbine engine of the turbofan type generally includes a forward fan and a booster or low pressure compressor, a middle core engine, and a low pressure turbine which powers the fan and booster or low pressure compressor. The core engine includes a high pressure compressor, a combustor and a high pressure turbine in a serial flow relationship. The high pressure compressor and high pressure turbine of the core engine are connected by a high pressure shaft. High pressure air from the high pressure compressor is mixed with fuel in the combustor and ignited to form a very hot high energy gas stream. The gas stream flows through the high pressure turbine, rotatably driving it and the high pressure shaft which, in turn, rotatably drives the high pressure compressor.
The gas stream leaving the high pressure turbine is expanded through a second or low pressure turbine. The low pressure turbine rotatably drives the fan and booster compressor via a low pressure shaft. The low pressure shaft extends through the high pressure rotor. Most of the thrust produced is generated by the fan. Marine or industrial gas turbine engines have low pressure turbines which power generators, ship propellers, pumps and other devices while turboprops engines use low pressure turbines to power propellers usually through a gearbox.
The high pressure turbine has a turbine nozzle including at least one row of circumferentially spaced apart airfoils or vanes radially extending between radially inner and outer bands. The vanes are usually hollow having an outer wall that is cooled with cooling air from the compressor. Hot gases flowing over the cooled turbine vane outer wall produces flow and thermal boundary layers along hot outer surfaces of the vane outer wall and end wall hot surfaces of the inner and outer bands over which the hot gases pass.
Film cooling is widely used in gas turbine hot components, such as combustor liners, turbine nozzle vanes and bands, turbine blades, turbine shrouds, and exhaust nozzles and exhaust nozzle liners such as those used for afterburning engines. Film cooling is used to inject cooler air through film cooling holes or slots to form an insulating layer or cooling film on the component hot surface and reduce the direct contact with the hot gases flowing over the component surface. The film cooling holes are typically angled in a downstream direction so that the cooling air is injected into the boundary layer along or as close as possible to the hot surface. The cooling film flow can mix with the hot gas and reduce its effectiveness as it flows in the downstream direction. The hot gas flowing over the component hot surface can lift the cooling film away from the hot surface and reduce the film cooling effectiveness. One method to improve the film attachment is to use a shaped film cooling hole having a downstream flare at an exit of the hole to reduce the angle between the film jet exiting the hole and the downstream hot surface. This method has been used in turbine airfoils where the gas velocity is relatively high. It is desirable to have another or additional apparatus and/or method that can enhance the cooling film attachment to the hot surface for good film cooling effectiveness.